Ramjet wing system for jet propelled aircraft



B. T. SALMON May 27, 1958 RAMJET WING SYSTEM FOR JET PROPELLED AIRCRAFT 3 Sheets-Sheet 1 Filed May 18,

JNVENToR.- @en /anwb 7 .5a/mon May 27, 1958 B. T. SALMON 2,836,379

RAMJET WING SYSTEM RoR JET RROPELLED AIRCRAFT Filed May 1s, 1954 s sheets-sheet 2 III www@

B. T. SALMON May 27, 1958 RAMJET WING SYSTEM FOR JET PROPELLED AIRCRAFT v 5 sheets-sheet :s

Filed May 18, 1954 United States Patent O RAMJET WING SYSTEM FOR JET PROPELLED AIRCRAFT Benjamin 'I'. Salmon, San Diego, Calif., assigner to General Dynamics Corporation, San Diego, Calif., a corporation of Delaware Application May 18, 1954, Serial No. 436,521

3 Claims. (Cl. 244-15) This invention relates to aircraft and more particularly to aircraft which are adapted for propulsion at supersonic velocities by ramjet engines.

Present day performance requirements for supersonic aircraft, including guided missiles, make desirable the development of a high-speed, self-propelled airborne vehicle characterized by extremely low aerodynamic drag and a high lift curve slope. As is known, an inlinitely thin hat plate is the lowest possible drag airfoil since its drag is only the shin friction drag of the two surfaces. lf it is assumed that such a plate is split in half and the surfaces moved apart a distance suflicient to provide space for a combustion system for propulsion, it will be apparent that a structure is provided whose external drag is unchanged as compared with the innitely thin plate, such external drag consisting only of the viscous or wetted surface drag plus whatever induced drag is present as a function of the angle of attack. The present invention provides a vehicle which is powered by what is in elfect such a split plate structure, since the structure is essentially a combustion system sandwiched between a pair of spaced apart flat surfaces or plates.

To produce thrust for propulsion of such a split plate structure, a momentum increase in the airflow is required and it thus becomes necessary to accelerate the air mass iiow passing between the two hat plate surfaces. It is weil known that this can be done most conveniently by the addition of heat derived from combustion of a suitable fuel injected into the air mass ow or air stream passing between the plates.

Since combustion can most effectively take place in air at low velocity and high pressure, a system is required to convert the high velocity, low pressure air encountered in the operation of rarnjet engines to the desired low velocity, high pressure state. As will 'oe described, an isentropic Oswatitsch type ofV supersonic diffuser is preferably utilized in the present invention to achieve this conversion.

To obtain suitable and efficient combustion, a burner system is required which is adapted for maintaining con tinuity of combustion, the system desirably including a high intensity pilot flame together with means for intimately mixing the fuel and air. Further, the burner system should be compact in size and capable of eiect ing high combustion efficiencies while utilizing the least possible length of air stream through a wide range of pressures at fairly high internal subsonic air velocities and at various fuel ratios ranging from stoichiometric to very lean mixtures. A unique can combustor with fuel injection upstream of the can is utilized in the present invention to attain these desired characteristics.

From the statements thus far made it is apparent that the greatest aerodynamic drag of the vehicle or missile will result by virtue of the system provided for burning the fuel in the internal airstream. Since this drag is not logically an aerodynamic drag but is rather a drag chargeable to thecombustion process, it is evident thatV a dimensionally thick structure is provided which is aerodynamically thin, since it possesses substantially only the true aerodynamic drag of the infinitely thin plate above described together with certain additional drag which is caused by the shape into which the leading edges or lips of the spaced plates must be formed in order to derive optimum efficiency from the supersonic diffuser.

The split plate structure of the present invention preferably forms both sustaining and propulsion means for an aircraft or missile, although it will be apparent that the structure may also be separately attached to a couventional aircraft such as at the wingtip, for example, to thereby provide propulsion means which are separate from the conventional wings or other sustaining structure. Thus, in the preferred embodiment of the present invention, one of the basic and unique features is the splitting of a wing along the chordplane into two coniponents, which components are separated by ducts and combustion chambers 'n which high temperatures and pressures are developed. As is evident, such a split wing structure presents unique problems relating to the interconnection of the split parts of the wing whereby the structure can eifectively resist the internal pressures and transfer the air or aerodynamic loads to the aircraft body or fuselage. ln addition,` such interconnection should be such as to provide minimum ,bending deformation, minimum torsional rotation, minimum obstruction to internal airflow, and good wall smoothness in the air or duct passages.

Accordingly, it is a principal object of the present invention to provide an improved, self-propelled vehicle capable of traveling at supersonic speeds.

Another object of the present invention is the provision of an improved airborne vehicle driven by a ramjet engine and characterized by low aerodynamic drag and a high lift curve slope.

A further object of the invention lies in providing a unique ramjet propelled aircraft embodying an improved form of supersonic diffuser.

Another object of the invention is'the provision of an improved form of aircraft in which there are provided ramjet propulsion means so constructed and integrated with the aircraft as to constitute a power plant which is compact in size and efficient in operation at its design speeds and altitudes.

An additional object of the invention is to provide for an aircraft a ramjet propulsion means of a configuration which not only serves to propel but also serves to sustain the aircraft in flight.

Another object of the invention is the provision of a rarnjet driven airborne vehicle which is characterized in part by a burner system and structure adapted for effecting ellicient and continuous combustion for extended periods without signicant localized deformation through overheating.

A further object of the invention is to provide an imroved ramjet propelled airborne vehicle which is charcterized by a construction and operation which contribute to improved fuel economy, extended range of operation, low vehicle weight, high capacity for payload, and minimum size of launching booster if such is required.

A still further object of the invention lies in the provision of a ramjetpropelled airborne vehicle which is simple and inexpensive to manufacture and thus adapted for use as an expendable missile.

@ther objects and features of the present invention will be readily apparent to those skilled in the art from the following specification and appended drawings wherein is illustrated a preferred form of the invention, and in which: v

'Figure 1 is a perspectivev view' of a4 guided missile embodying ramjet propulsion means in accordance with the presentinve'ntion; i

' Figure 2 is a perspective detail view of a portion of the Yport wing panel 4ot the guided missile, portions being j omitted for clarity;

Figure 3 Vis a Vperspective' Vdetailfview illustrating Va verticals .23 each being'secured to the forward end of one typical connection between upper and lower wing surface panels and a chordwise rib panel ofthe wing; l Y

Figure-4 is a chordwise sectional View of the port Wing panel taken along line IV-IV of Figure 6; l Figure 5 is a frontY elevational view of the port Wing panel, taken along line V-'V of Figure 4, the forward Y the invention which is ada'ptcdrfor use asa pilotless aircraft or guided missile, although it will be apparent that the wing of the missile may Valsoeasiiy be adaptedV for Y otherV uses, such as for use as an external ramjet pro-Y naamw. 'K

of rib panels V22, and forming a forward extension thereof, while intermediate verticals 24 are each located approximately midway between a pair of mainverticals 23 and Yserve as an auxiliary spacing and'support meansrfor spar caps i9. Spacing and support means for rear spar capsV 21 are provided by a plurality of rearward verticals 30 which, like main verticals 23, are also uniformly spaced in a spanwise direction, and arerrigidly connected, as by i welding, to the rearward lendsiof panels 22 and, at their Y upper. and lower ends, to sparcaps 21 to thereby form, with a corresponding number of nozzle verticals 25, rear- V i ward extensions of panels 22. Y it is noted that there are no intermediate nozzle verticals corresponding to the described intermediate verticals 2 4,V this being so for the reason that the strength characteristics of wing panel 15 have been found to be suiciently good that the omissionof intermediate verticals yis permissible and desirable since it will reduce internal flow restriction.V

pulsion means for a conventional piloted aircraft byisuit- Y able attachmcnt'thereto( Forconvenienceof description the particular embodiment'illtistrated hereinis sometimes referred to as aV split` Wing ramjet, this name being ap- V'propriateby .virtue of the double function or" the wing as Ya sustaining and a propelling unit. Y Y

' The Vaircraft or missile, designated inits'entiretyby ,the numeral 11, comprises generally a fuselage 12, yempennage13, and arwing i4 which latter constitutes the primary airfoil VVor sustaining means lfor the aircraft.. 1Wing 14, which also constitutes the ramjet propulsion rn'eansforY the aircraft V11,V includes ay port wingfbody, housing, or panel 15 Yand a starboard wing body, housing or panel 65' Since wing panels 15 'and 15 are identical in Vtheelements'in-` Veluded in their construction, the detailed descriptionhereinafter made will `be primarilymdirected to port wing V.panelli-. *Y Y I Y Wing panell, as shown, corr'rprises,` generally, a pair of identical, substantially uniformly spaced apart outerV members or wing surface panelsjt?V which are connected.. at. their'outboard ends'by a wing tip 1S whose streamline shape, as illustrated,A tends toi reduce the aerodynamic drag offwing panel 15.VVV Y Forming a forward portion or extension o/f'each ofthe upperfandlower wing. s'urfacepa'nels i7 is a spar cap or Vdiffuser lip ,19,Yand' 'forming' area'rward extension ofV each wingsurface panel 17isa spar capV 2i. .Spar caps 19and21-proivide the principal spam-vise strength for Vwing 14, earchv of the forward pairV ofspar caps 19 being!` rigidly securedY to Vthe Vforwardy edges of its associated I surface panel 17, Yand each ofV the Vpair( of rearward spar ribs 22 in the space'or bay defined therebetween.` Di-` A burner, combustor or can 26 is located between surface panels 17, as best shown in Figures 2 and 4, and is Ypreferably sectionalized, Vas will be more particularly described hereinafter, to t between each pair of 'adjacent Viding can 26 into a plurality of burner or can sections as above mentioned is done merely in order to simplify` and facilitate assembly of wing 14 during theQconstruction or Yservicing thereof.

Can 26k serves as a ame holder and provides a means i Y for effecting combustion ofthe airfuel mixture which .'ili berfed thereinto, and is'serv`ced by a conventionalY fuel supply and vinjectionsystem which is illustrated in sufcient detail for'iinderstanding thereofand'fis designated generally by the numeral 27. Located forwarder upstream of can 26 is Yaninlet wedge or diffuser 28 which servesto effect supersonic -isentr'opic external compression and subsequentsubsonic internal expansion of ramair l entering can 26, Vancllo'ca'ted. at the' rearward end of can 26 is anozzlev section 29 of preferabilyconvergent-diver-`V gent conliguration (Figure 4) vwhich is lremovably :se-V

l cured to'rear spar capsr21 to thereby Ydefine ari exit for combustion' gases escaping or spouting fromV can` 26.V 'Y Diffuser 2S terminates at fuselagelZ andjs disposed and supported betweenthe pair of front spar caps"19"by Y main verticals 23 and intermediate verticals 24, diffuser 7.29 therebyV providing an vadditional or auxiliary beamer` spar for'spanwise strength and reduced deflection of Wing Vcaps 21A being ArigidlyY secured to the rearwardV edges, j'

of its associated surface paneli'i7. It is here noted Ythat the pairs of spar caps,a VandiZl are not only continuous along `Vtheforward andrrearvward portions, respectively, Vof w'ing'lbiit also desirably run' through, and are suitablyV affixed t0, fuselage 2 to thereby'providespanwise strengthV and' rigidity for vvin'gjl'sii.V All' Iother' Vportions'o wing 14VV Y terminate or are discontinuousv atjtne side of fuselage 12. Y The means and method' which may be utilized to afhx or Ysecure Vwing' M toffuselageZ are not important'to the r presentinvention, any conventionalimeans `fOrgidy maintaining wing 14 Yin'position being satisfactory.

j Wing surface panels 17', asbest shownin Figures 2 6,/ areY securely maintained ingspaced apart Y pai'aliel relationship 'by a pluralityof;niainribs or rib panels 22 which are substantiallyv uniformly spaced infa spanwise l l direction, andextend in a cho'dwisedirection between the fp air of V*surface panels 17j toth'ereby" provide additional VVAchordwise strength VV.for wing panel l5, and also to form,V jwith .wing surface panels Jr?, ansop'en endedV enclosure 5 jiliS-Q Similarlypthepair YVof front'spar' c'apsrl9V are Vspacectapartar-idsupported' byrra'pluraligtyof main verticals 23and aplur 'ty V'ofintermediate verticals:24;main

l fable-' degree'of dimensional changeV infpanelA lif; .K A- i jEach wing' surfaceV panel'.17.possessesrcomparatively high 'strength-weight ratio byvirtue'lofits:uniqiiecornaj j Ypanel 15. Similarly, nozzle section V29,"by virtue of'its spanwise disposition coextensive withk rear spar `caps 21, i

provides further spanwise strengthY forA Wing panel 15'.

VHere'inabove,` aV description has been Yafforded of the Vrnain elements of the inventions VandV their interrelationshipVV and nowa more detailed individual description of thesevarious portions and components of Vthe v present invention and their purposes will be made;

The Vprimary structural function of `wing surfaceV or skin panels 17 is to transport the .operating ram pressufeV orbursting pressure developed within wingl panel Y:l5 to` Y the pluralityA of rib panels 22'where Vit iscanceled ,orvr equalized by the loadbearing capacity of surface'panels'vv 22. Also, wing surface panels i7 transmit the aerody`V namic lift through rib'panels 22 to V,spar caps 19 and 21 and thence to'fuselage 12 of the aircraftyand further sur,

face panels'l' Vtend to equalizedeectionsof spars V19 and ZL In these circumstances it'hasgbeenlfoundithat the construction of'each'wing surface panel'vr 17 'and rib panel'22 in the present invention provides Yhigh strength.VY

' and isY advantageousV inV that. the described structural: functions Yare effectedV with ja low degree of dirnen-V sional changeinwinglpanel 15 Vwhich hasbeet'rfoundA to;be `important sincegtestschave indicatedthe elliciency v otfdius'er 28 is sensitivato'any pronouncedorappreci-- posite construction. This construction is effected, as best illustrated in Figures 2 and 3, by the welded securement of a plurality of vertical elements 31 between a pair of skins or sheets 32, which form the outside and inside surfaces of each wing surface panel 17. Elements 31 are uniformly spaced chordwise along sheets 32, and are continuous in a spanwise direction. This composite construction is identical in each of the pair of upper and lower wing surface panels t.

The construction of rib panels 22, which space apart wing surface panels i7, is substantially identical to that of wing surface panels 17, each rib panel 22 including a pair of vertically disposed rib skins 33 which extend in a chordwise direction and are securely maintained in spaced apart relation by a plurality of elements 34 disposed normal to rib skins 33 and preferably welded thereto.

Each rib panel 22 is connected at its upper and lower edges to a chord-wise extending stub rib 35 which in turn is welded to wing surface panel 17. Each stub rib 35 includes a plurality of stub elements 36, similar to elements 34 of rib 22, which are welded normal to and between a pair of spaced stub skins 37, similar to skins 33 of rib 22, whereby upon assembly the components of each stub rib 35 become in effect extensions of the components or rib panel 22, facilitating assembly of wing surface panels 17 and rib panels 22 by virtue of the comparatively easy accessibility for welding of the butted joints, as at 38. Thus, there is provided an integral wing surface panel 17 of composite construction which is characterized by high strength and good dimensional rigidity. it is to be noted, however, that wing surface panels i7 are preferably discontinuous through fuselage 12, the principal spanwise strength of wing 14 being provided by spar caps 19 and 21, as previously stated.

Front spar caps 19 together with intermediate verticals 24 and main verticals 23 act generally as a spar structure which is similar to the well known Vierendiel truss, spar diagonals being preferably omitted from the spar structure for various reasons including the more desirable ram air flow properties thereby achieved. Each of the front spar caps i9 of this spar structure comprises a formed member embodying, as best shown in Figure 4, rearwardly extending upper and lower flanges 39 which are each provided with a spanwise extending step 40 to which sheets 32 of wing surface panels' i7 are welded. In addition, front spar caps 19 are provided with a lightening channel 4l extending in a spanwise direction and which is closed by a plurality of spar plates 41a welded in position as illustrated. Each of the pair of spar caps 19 is curved inwardly at the forward edge thereof so as to provide a curved lip portion which in combination with the forward portion of diffuser 23 serves to achieve improved efficiency in converting the velocity energy of high speed ram air entering wing panel 15 to pressure energy for improved combustion.

As previously mentioned, main verticals 23 and intermediate verticals 24 space apart front spar caps 19, the upper and lower ends of main verticals 23 each being rigidly secured to a separate connecting tab d2. One tab 42 is welded between the upper edges of main vertical 23 and the base of the formed lightening channel 4i in the middle portion of one of spar caps i9, while the other tab 42 is welded between the lower edges of vertical 23 and this formed lightening channel di. It is noted that spar plates 41a are welded at their ends to tabs 42 to further strengthen the connection of main verticals 23 to spar caps i9. in a similar manner intermediate verti-y cals 24 are rigidly co-nnected, preferabiy by welding, to spar caps 19 by a pair of tabs 43 which are welded to spar plates 4in1 in a manner like that just described in connection with verticals 23.

Each main vertical 23, as best illustrated in Figure 6, is constructed of a vertical plate 44 and a vertical plate 45 whose forward edges are shaped to a point and Weldably joined, and whose rearward edges are channeled,

as at 46, to weldably receive rib skins 33. In addition, a spacer 47 is weldably disposed between rib plate 44 and rib plate 45 to maintain and strengthen the spaced relation thereof. Similarly, intermediate verticals 24 each comprise a pair of vertically disposed intermediate plates 48 and 49 whose forward and rearward edges are shaped to define substantially knife-like edges, the rear- Ward edge being notched at 51, Figures 2 and 6, for receiving in welded connection an igniter fitting 52. In this manner, the combination of spar caps 19, main verticals 23, and intermediate verticals 24 provide a lightweight composite structure of comparatively high strength which is adapted for effectively resisting the combined forces of internal pressure and external aerodynamic pressure in a manner structurally similar to the conventional Vierendiel truss.

As described previously, diffuser 2S serves as an auxiliary beam for the Vierendiel truss structure just described, and also serves to efect isentropic compression of entering ram air and suppression of shock waves or pressure and velocity discontinuities to thereby utilize a maximum proportion of the kinetic energy of the entering ram air, as is well known in the art. Diffuser 2S preferably possesses a cross section similar to the cross section of the Oswatitsc'h type diffuser, which type of diffuser is described in Elements of Aerodynamics of Supersonic Flows, by Antonio Ferri (The MacMillan Company, 1949), at pages 193-195. The type of diffuser used in the present invention is a version of the Oswatitsch multiple oblique shock type, utilizing a continuously cur-ved concave control body profile, instead of a multiple wedge, for external diffusion. The diffuser is designed to produce two dimensional supersonic external isentropic compression of the air ow along the cont-inuous concave central body with an accompanying reduction in Mach number to some value above 1.0 at the diffuser lip. A normal shock occurs at the lip and the air is expanded subsonically to the desired value to permit satisfactory combustion. In order to obtain maximum mass flow at the designed Mach number the shock wave from the forward edge of the diffuser 28 and the compression waves from the exterior surfaces of the diffuser are all focused at the diffuser lip i9. No internal supersonic compression has been allowed in the present invention although, if desired, such may be provided for. In addition, a half-Oswatitsch type diffuser may be employed as a variation, if desired, of the full-Oswatitsch' type herein preferred and described.

It is particularly noted that the present diffuser is mainly two dimensional in configuration, ow occurring over and under the central diffuser body, or simply over in the 'case of a half-Oswatitsch type diffuser body, as contrasted with three dimensional flow in which flow occurs, for example, about a body of revolution which presents a circular cross section to the fluid flow.

As illustrated, diffuser 2S is characterized by a spanwise extending knife edge wedge portion, or forward portion 53 which flares or diverges smoothly rearwardly to join other structure of diffuser 28 in the region of maximum lcross section thereof, diffuser 28 thence reducing in cross section to a junction with can 26. At its rearward portion knife edge 53 is stepped, as best shown in Figure 4, for weldably accepting the forward edges of a pair of smooth surfaced forward diffuser sheets 54, as illustrated, the shape -of sheets 54 serving to dene the configuration of diffuser 23 from the forward knife edge 53 to appro-xi- -mately the region of maximum cross section of diffuser 28. Each of these forward diffuser sheets 54 is preferably removably secured at its rearward edge by any suitable fastening means sucn as by machine screws 5b or the like, within a step provided in a respective one of a pair of vertically spaced apart and spanwise extending lip spars 55. The removable connection of sheets 54 to lip spars 55 permits easy disconnection thereof for inspectiongor servicing of fuel supply and injection system 2f! l and usual associated ignition wiring. Spara `SSart Yalso provided, as shown in Figure 4, with stepsatytlgteirY fuser 28. vThese sheets 56 are disposed in Vopposed pairs between and welded to each main vertical 2,3 Yaud-its adjoining interrnediate vertical 24, or in the inboard region' l Vof wing section to fuselage 12.. Further rigidifying the integration of diuser 2'with other structure of wing panel 15 istlie welded attachment of .lip spars 5S within suitable Vnotches 55a provided in the forward edges of main verticals 23 andintermediate verticals- 24. In ad-Y n dition, `the outer surfaces of diffuser 2S are preferably made smooth Yin order to achieve -good aerodynamic en YDiffuser 2S is Vinternally braced or strengthened by a Y series of members which are vertically disposed'and eX-V tend'in a chordwise direction, these members including a plurality of wedge ribs 57 which are uniformly spaced in a spanwise ,direction Vwithin the space between the pair of forward diff-,user sheets 54, and suitably weldedV thereto.Y Strengthening of the rearward portioniof diffuser 2S is provided by a plurality of after wedge ribs SS'whicli are spaced, `asbest illustrated in Figure 6, in a spa'nwise direction,Y two of ribs 58 preferably being'disposedV bef .8 i vfariallearea exit type nozzlemay well justify its userid many cases where the=increased complexity therebypr duced lis`V not objectionable. s, AlthoughV some form Yof nozzle section isa necessary part of the combination making fup port vwing -panel '15, the particular forn'irofY nozzle Ysection 29 illustrated in the drawings Ydoes. not forman essential partof fthe present invention but instead is merely one'form of nozzle section 29 which has :beenVV found to give suitable results. k n Nozzle sectionV 29 hereinV utilized comprises ya plurality of nozzle yerticals 2S which are welded Vat top and bottorn` in'spaced, spanwise relation, between the inner surfaces of a pair of nozzle fairings 67, as best illustrated in' FiguresAZ and 4. VThe pair of nozzle fairings 67 serve Y to denne a restricted exit or throat for the `sonic or s nper-jy smid-expulsion of combustion and other gases leaving'V can Nozzle section 2% is geuerallyof lightweight composite construction, the various components being. weldedV together-to form an integral unit. Thus nozzleVK verticale 2.5 each comprise a pair of nozzle Vplates 68 .j which are spaced apart at their forward edges, closed at their rearward edges to form a rearward .knife edge, and" rearward end portion to rweldably receive a guidebushing Y 59'for aligningand securing can 26 on assembly of wing.

1'4. As' illustrated, the Vremaining after wedgeribs 5? are not notched orc-ut away, but are instead continuous from yone rearward diduser sheet 56 to the opposed'sheet 576. f

In'addition, in Vorder to close .the rearward open end of diffuser 28V and for addedstrength a vertically disposed rearwardA bulkhead 61 is welded t-o and extends ina spanwise direction between each main vertical'ZS and the adjacent igniter-tting 52, eachrbulkhead 61 being provided with appropriate cutouts or openings for tionof Vguidel bushings 59 and igniter fitting Y52.' Bulle'- heads 61 are also welded tocontiguous portionsof wedge ribs 58. Further internal bracing for'diffuser 2S is pro- Vvided by. a series of intermediate frames V63Y which 'are` YThe .forward portion of each of thev pair shaped at their upper and lower edges to conform withV Y theinner surface of fairing 67".V In addition, theV forward edges `of nozzle plates 68 are adapted 4to removably mate withwcorrespon'ding vertical channels provided in rear- -wardvertical's Tunas illustrated best in Figure V6.V Each nozzle"fairing 67 is `also composite in construction, com-s Y p'ri'sing an o'uterlplateV 69l and, an inner plate 72-wbcl2tf` are'forrned, as illustrated, to dene .thefnoz'zle throat or. eXit ofthenn'ozzle section 29. Y` e i @f toner' plates 69 is formed toY provide a forwardly extendingrledge which is. removablytsecured, by suitable fastening.

means, .such as machine screws 7i, Within a complement-1 -tary channelprovided in the Vrearward portion. of la re spective rear spar cap 2l; with this arrangementnozzle",

t section 2? mayfbe removed toV obtain access tothe inf terior of wing panel i5. [Y

n in addition, the forward por-` .tron of each .outer plate@ is provided with rearwardly dii@ Y rected portionrwhichY is weldedto the forward'portion'of a accommode# vertically v'disposed.and uniformly spaced spanwise Vbe.. ,Y

V'tweenV opposing after wedge Vribs 5S, being welded to' VribsS to provide increased strength forY resisting aero-1,. ,-dynami'cloads exertedupon diffuser 28 by ram air enter- 'jing wing panel l5.V Y

The interiorY of diffuser Vf'is V'hollow for light weight pose of reducing aerodynamic drag by presenting a clean, streamlinepath'for entering ramV air. Thus, for'exarnple, al main fuel line 64.and a pilot` fuel line 65 Yoffuel system 27 areldisposed within diffuser Y2S through a plurality of Discharge 'nozzlef-or Ynozzle sectiong29. is Vof the .icon- Y' f'vergent-divergent type and serves as a-sonicf and super# andV to'enable certaincomponents ofthe pres-entv inven- Y .tion, including. fuel supply and'injection system Y27,Y toibe .substantially ycompletelyhoused therewithin for thepurf Y respective inner plate72, The rearward edges of plates;Y

69.V and V72 are welded together to form a pair ofes'paced" plurality of substantially uniformly spaced fairing ribs l' 73 `are provided to r'igidify nozzle fairings 67, fairing ribs 7,3 embodyingV openings, as illustrated, for desirable weight reduction. Y Y in' assembling nozzlejsection 29 to Ythe otherNsti-ucl' tureof wingpanel l5 it isY desirable, as above stated,to

employ readily removable fastening means, such V'as screws 7i, to permitA easy detachment ofVY nozzle Vsection '29 in order to provide ready access tothe interior of wing panel i5 forVV various reasons, such as to'enablethe'n insertion or removal ofjone'ortmore sections--ofcan'26;. Y"As previouslydesc'rioed, Caril is sectionalized andV comprises apluralityof sections 74, there being a Vsegr,l

Y.tion for each bay of port wing paneltl'g. Thus, in the bay formed between thejpair of mainV verticals .Bf-,illus-Y trated in Figure 2,'thereV is'l'provided :a typical, can sec-'f t t knife edges extending in a spanwise direction; Further, *a5 j VY tion 74,'it being understood that additionalcansectios fr it are utilized in the other bays` of wingV panel-15 Ybrutare Y "j portion `*of diffuser ,231mY slightly beyond.thef forward Vor Y upstrearn'portion of nozzle section 29g/and, Vasis"best11 sonic propulsive Vnozzle for thefejectionlofheated air and combustion gasesV Vgenerated by Ythejombustion processf .whereby there is derived the thrust forp'ropulsion .of misf Y sile ligas' is .well known toV thoseskilledint'he art..V .Noz-`Y Y VV'zie section has a fixed area exitort'hroat in4 the .prese f ent invention, ,but it iis evi-dent that nozzle sectioni59mayv .i

t 'Y beproyidedrwithavariable area exit'orthroat'if desired.

j for1 'in`creased eftciencflr.-Y As Varrna'tterl of fact .the *in-t. Ycreased'efficiencyV which rnayjbe` achieved byY utilizingai omittedin'Figurel. for clarity: ofillnstrationt Can sec-V t tion 7 4i, isrmainly two-dimensional Vin function,` muchilike f f diffuser 2S, and Yextends rearwardly Vfrornthe downstream apparent'ffrom Figure 4, can ysection 74,', diffuserzlir'nd nozzle k,section 29 are positioned vwith their longitudinali Vartes/.in.alignment;,YV 1 'Y Can section 7,4`is' provided Vwith aftotribus/tion:chamberVv 75g',V which'isdeiinedbya pair'of vertically ,spacediburher ,fr

plates '16. aflaieh'fforni the-top and bottom" wailsoflthe chamber and Whichare interconnected by an integral j Y forward-.burner Wall 77. s The side chamber walls are' l',

aorded by a pair of burner side wall plates 73 which are welded to the lateral edges of plates 76 and serve to space apart and support ;'lates 75. Side Wall plates 78 and rib panels 22 are preferably provided with an opening or plurality of openings (not shown) to eiect communication between the various bays or sections of Wing panel whereby substantially uniform combustion pressures will be achieved throughout these bays. The location of such holes is not critical and need only be such as to permit the function described.

lt is apparent from the drawings that the configuration and association of burner plates 76, forward wall 77 and side wall plates 73 are such as to form a box-like burner or combustor structure which in external appearance, as best illustrated in Figure 4, flares or divergcs rearwardly and is open at its rearward or downstream end. This box-like burner structure which includes the combustion chamber or fuel-air mixture zone 75, is supported between wing surface panels 17 by a plurality of runners or support or hanger members Si, which are equally spaced spanwise over the surfaces of plates 76, each hanger member di extending from a maximum height in the area of forward burner wall 77 to a minimum height at the downstream end of can section 74. As illustrated, each hanger member S1 includes at its outward edges a flange 82. which is adapted for sliding engagement with a corresponding hanger track, as at 83 in Figure 2, the can section 74 for that bay having been omitted to more clearly illustrate the position and shape of track 53. Thus, with nozzle section 29 detached from the wing surface panels 17, can section 74 may be slidably inserted between wing surface panels 17, each flange S2 slidably cooperating with its respective track 83. When inserted into position, can section 7d extends in a span- Wise direction between' adjacent rib panels 22 and is positioned slightly spaced therefrom to aord passage for cooling air exteriorly of the side walls 7S of the combustion chamber '75. Further, the runners or hanger member Si will in cooperation with the interior surfaces of the upper and lower wing surface panels 17 define a plurality of passages or ducts for flowing the stream of incoming ram air above and below the top and bottom walls 7 6 of combustion chamber 75. it is noted that the air passages or ducts defined between burner plates 76, which form the top and bottom walls of combustion chamber 75, and the interior of wing surface panels 17 rearwardly converge to a minimum section at the rear of can section 74, and likewise the passages or ducts defined at each side of can section 7d by side wall plates 7S of can section 74 and by adjacent rib panels 22 me rearwardly converging. rille converging ducts or passages thus formed are designed to accommodate streams of air for cooling all sides of can section 74 during the operation of aircraft ll.

Can section 74 is rigidly secured in position by a pair of studs 85 which are integral with forward burner wall 77 and are disposed through guide bushings 59 of diffuser 2S, a pair of suitable lock nuts S6 being provided to complete this connection. ln addition, an opening is provided in burner Wall 77 so that upon connection of can section 74.- to bulkhead el, the rearward portion of igniter fitting 52 is open to the interior or combustion chamber 75 of can section 74 whereby fuel and ignition means may be provided for the pilot ilame, as will be seen.

Burner plates 76 of combustion chamber 75 are provided with a plurality of perforations or holes S7, which, as illustrated, are progressively larger toward the rearward end of the plates 76; further, various of the holes 87 at the upstream end are, as shown, provided with a scoop 8S formed by outwardly dishing the surfaces of plates 76. At the downstream end there are provided a plurality of openings or louvers 89 which by virtue of their shape serve to bring cooling air internally at the rearward or shroud portions of can section 74. These sastre tft E louvers are located at the upper and lower wall plates 76 of can section 74, and also if desired may be pfovided atV the side walls 7S for additional cooling. Thus, a stream of air flowing over the surfaces of plates 76 will be initially diverted in part by scoops Sinto combustion chamber 75 and the remainder of the air will then proceed downstream and a substantial amount will pass through holes 87 and louvers 89. As will be seen, this ilow of air serves not only to provide air for combusu'on purposes but also serves to provide a cooling film or layer air over the inner surfaces of burner plates 7d to reduce the temperature thereof.

Fuel supply and injection system 27 serves to provide the fuel for combustion, and includes a plurality of rearwardly extending fuel nozzles or fuel injection tubes 31 located in abutment with the outward surfaces of plates 76 and rearward diffuser sheets 56. Fuel tubes l are disposed through suitable openings provided therefor in sheets 56, and each connects through a usual restrictive orifice section (not shown) to main fuel line 64. In addition, a pilot fuel tube 92 from pilot fuel line 65 is connected to a suitable fuel nozzle 93 which is threadably secured within nozzle tting 52. Main fuel line 64 and pilot fuel line 65 are connected to any suitable fuel supply means, such as a pump or pressu'rized tank (not shown) located within fuselage 12. it is to be understood that any usual and conventional fuel supply system may be utilized in connection with the present invention, it being important only that an adequate supply of fuel at the proper pressure be available for combustion within chamber 75 of can section 74.

Nozzle 93 serves to supply only a small or pilot amount of fuel for the maintenance of a pilot or ignition flame in the region just downstream of burner wall 77 where the velocities of entering air are comparatively low. Fuel tubes 91, on the other hand, supply the fuel for the principal combustion occurring in chamber 75, the fuel being intimately mixed with incoming ram air by the high degree of turbulence created by the passage of ram air through scoops 8S and holes 87 and the consequent irregular air ow within chamber 75.

The ignition system of the present invention is not shown in any great detail inasmuch `as the particular form of the ignition system used is not important to the present invention, it being sufficient that a system be used which is adapted for creating a spark or other ignition means in the forward or pilot llame region of com'- bustion chamber 75. Thus, for example, .a spark or glow plug 94 may be threadably secured in igniter fitting S2, as best illustrated in Figure 5, plug 94 projecting into the forward portion of combustion chamber- 75 to thereby provide means for igniting the pilot flame and maintaining steady burning. The energy source (not shown) for actuating spark plug 94 may be located in fuselage i2 and comprised of batteries, ignition coils, and condensers to thereby effect a capacitive discharge type of ignition characterized by a high intensity sparlt.

The inboard section, designated generally by the numeral 95, adjacent fuselage l2 differs from the typical bay of wing panel l5 mainly in that a burner can section 74 is omitted therefrom to allow a comparatively free flow of air therethrough. Thus, as best illustrated in Figures 5 and 6, the inboard tip of diffuser 28 is made concave, in elevational cross section, in order to fit against fuselage i2, and wing surface panels 17 and nozzle section 29 are extended to form a part of the inboard section. The omission of a can section therefrom is for the reason that control is thereby had of l-ow energy boundary layer air flowing about fuselage 12, and also to obtain a cooling of the area adjacent fuselage 12.

The outboard tip section formed by wing tip 1S serves to reduce drag and is generally comprised on continuations of the components of wing panel i5, the configuration of the wing tip section components mainly differing to the K extent offtheir semifcircular shape and the three-dimensional ow thus'Y occurring about'their surfaces. As

e respects wing tip 'section i8, a cansection modied to be accommodated within theconiiguration of winggtip section 1S maybe fitted therein if desired. Details of construction of such m'odied can section are not here shown Y ff'hecomponents of`wing tipV V1.55 may be secured in posi- V'tionrto'its wing panel section Y15 by any suitable means such'as, for example, Welding, or by machine screwsV or the like if arremovable attachment is desired.

M A pair of Ycontrol'surfaces or Aiilppers 96 mayl'be'fat# tache'd towing tips'l, respectively, and secured Vthereto in anysuitableV mannerand operatively coupled Yto usual andA conventional control means or apparatus located in fuselage ,12. Control surfaces 95'se'rve as lateral'rcontrol forY aircraft lland are not shown in detail forthe reason Vthat Vthey do'not form an'` essential part of the presentV invention, Vbeing only suggestiveof aconvehtional control means. Y

A suitable corrosionk and temperature resistant metal 'Y is employed in the present invention. A preferredfmetal isV l7'-7PH stainless steel which is commerciallyfavailable Y j from theAmerican RollingiMills Corporatiom Welding of this material may be by either of both the electrieresistance weidingmethod or metalarc welding method, with 1 Vfor Ywithout filler rod but preferably' inert gas shielded.` It ia important, however, that with any materialsand 'attachrnent methods'used that the resulting structure be adaptedto `withstand the fairly high temperatures gen- Y erate'd'in the use of the apparatus of the present yinven- Yf tion.V VIt is specifically understood that `the ,particularmaterial and fastening methods mentioned are merely prcferred yand that the yuse of suitable substitutes are {contemplated. Any material employed desirably should exhibit,Y inV addition to good high Vtemperature characteris tics, Varphigh Youngs modulus ofV elasticity, at elevatedVV temperature',V for stiffness; reasonable elongation for i. formability; a low coeicient ofUther/ma'l expansion and easeof welding.'V Y v Y Y Y VVThe' present invention isV designed to function with loptiinum results athigh. altitudes'of between approxi- Vmately Vv30,900 to 1001000 feet, andratV a flight Mach number'of approximately 3.0. it isn'particularly Vto'be noted however that these'are merely desirable limitsof operation, and that no limitationrofY Ythe application ofY the present invention is intended.

Tne'oreration andV ur ose of `nin anel 15 in corr-V Y :l i

Vjunction.with aircraft llhas been generally described hereinabove, butit is believedV desirable to Vsu'rn'n'i'arize the Voperation or Ytypical functioning of van V aircraft. 1l'

when equipped with the unique wing i4 of the presentV invention. Thus', aircraft' i1 is `first Vbrought', by Yair Vor .Y ground launching, asidesired, to a speed suitable forthe selfpropelle-:l operation ofthe `rarnjetl type of'j'et Ypropulsion,Y .thisV launching 'generallyA being accomplished'V throughthe usefof expendable ,rocket type boosters, as

'Y is well known in the art: '.In'this connection it is to beV noted thatrthe present'invention possesses advantages over assasr iZr Y altitude have been attained.l Further, in addition to the Y aerodynamic lift for missile 11 which is achieved by virtue of externalairflow over external surfaces of vwing 14, there is additional lift derived from `the internal Vtlow of air .through Wingle. Y l Y As isrwell known, self-supporting,combustion in ram--V jet engine propulsionV means, such Yas isrembodied in mis.- sile 11, does not occur Auntil the speed of the ramjet rela-V tive to the outside airstream is sufficient that the ram effect of airentering the ramjet elfects ak compression ratio great enough to support combustion. VThus,-in the present;` invention the velocityrenergy of the high velocityfair stream entering wing 14 is converted byY diffuser 28,*in

t combination withditfuser lips i9, to a much lower velocitV airstream of higher ressure ener at the -u stream- VVend of can 2e. Y This locai slowingdown of theairstream" produces' a condition desirable for effecting a steady igni` Y tionV zone within can 26.

Fuel is then supplied ztllrough pilotV fuel vtubes 92 Ato fuel Ynozzles 93V and is finely dispersed or sprayedrwithin the upstream end of can 25 where it Vis ignitedby glow 4plugs 94, creating a steadyY ignition zone which servesv as a pilot zone or flame for the main Vcombustion occurof the main .fuel air` mixture.

Fuel from main fuel line 641s carried bythe plurality Y ofjfuel Vtubes 91 to a position slightly upstream Vof scoops 88 of can 26,'as illustrated. VUpon leaving tubes 91, this*F fuel is finely disbursed bythe turbulence oftheairstream4 as'it 'enters scoopsS and, subsequently, holes v87, and., the resultingfuel-air mixture incombustion chamber75 YV Of can 26 is ignited bythe pilot llame just described. TheV t turbulence of theV airstream at this point, created by Vscoops 88 and holes', also asists in the mixture of burn- V l ing with unburned gases Vto hasten the combustion process Vin combustion chamber '75 wherein the combustion proc,-V

ess is'carried substantially to completion.

A considerable quantity of what is'termed 'Y air isconsurned in the Vra'mjet combustion process, but a aircraft o'nmissilefli can itself assist inthe attainment f in the Afuel, avaiiable ,to missile idonee design speedlan'd That is, as .the booster means startsV Y missiie', the renier ofwing ld maybe utilized almost .p immediately, assun'i-n'glthe.eXistence'of self-supporting combustion; in order 'to thereby assistini-boosting Ymissile Sii' at lower, off-design speeds. and altitudesV and luntil 'g design speed and altitude `iis reached. lThus,-the size and k.1 'Y

great deal of the entering airstream is not psorconsumed, but instead is used as secondary air for coolingpur-y lposes. In connection, secondary air entering holes 87, andparticularly entering louvers 89 positioned Von can 26-where temperaturresare fairly high, provides a blanket',A or boundary layer of air which sweeps Vover the innersur- Vfaces 'ofV can 26 for VcoolingV thereof.` YAnother Ywell i known eect Vof ythis boundary layer` of Vcooling air is the scrubbing fromrthese inner walls of soot or carbonaceousV deposits which may be Yformed by the irnpingement of.- fuel against the hot structure defining chamber 75,@111` addition, this secondaryair ows in the ducts;adjacent'.

wing surface panels i7, and further, is mixed with, .or

dilutes,rtherhot gases in chamber '75 whereby an aceptable'.

' temperature Vdistribution is 'provided throughout Vthe structure of Wing 14. Y

Y 1 There has been describedhereinabove'a structure'ap'` .Y proximaitng the ideal, low drag split plate structure and a which forms bothsustaining and propulsion means `for anV aircraftrvehicle, and which 4is characterized by high l' sional airstream ow.V

Y While 'certain preferred'embodiments of Vthe .inventionV 4. 1 have been specifically disclosed, it isunderstood that the invention Yis.. not jlimi te""r 1t`neretoV as many variations 'will j be readily apparent to rthose skilledY Vin the art andthey .Y g

invention is to be given its broadest possible interpretation withirrthel 'terms of the'v'followingclaimsz. i y

What Iclainris:

l. nvan airbornevehicle, acrodynamicsustainingand V..propulsionmeans operatively associated `Ywith said velucie, saidrr'neansn comprising a pair of spanwise extendL ing panel members arranged in spaced superposed relationship, a spanwise extending air diffusing body located between and spaced from said pair of panel members at the forward edges thereof, said air diusing body having a forwardly extending wedge portion whereby it is adapted to eiect two dimensional supersonic external compression of the air flow, a spanwise extending combustion chamber located between said pair of panel members, said combustion chamber having apertured waLls forming a faired rearward continuation of said air difusing body and diverging rearwardly to form converging air passages with said pair of panel members, and nozzle means located between said pair of panel members and forming a rearward continuation of the walls of said combustion chamber.

2. ln an airborne vehicle, propulsion means forming a wing of said vehicle and comprising a pair of spanwise extending panel members arranged in spaced superposed relationship, each of said panel members comprising a forward spar, a rearward spar, a pair of metallic sheets connected between said forward spar and said rearward spar, and a plurality of transverse sandwich elements connected between said pair of sheets in a spanwise direction to thereby rigidity said panel member, a spanwise extending air diffusing body located between and spaced from said pair of panel members at the forward edges thereof, a spanwise extending combustion chamber located between said pair of panel members, said combustion chamber having apertured walls forming a faired rearward continuation of said air ditfusing body and diverging rearwardly to form converging air passages with said pair of panel members, and nozzle means located between said pair of panel members and forming a rearward continuation of the walls of said combustion chamber.

3. In an airborne vehicle, propulsion means forming a wing of said vehicle and operatively associated with said vehicle, said means comprising a pair of spanwise extending panel members arranged in spaced superposed relationship, a spanwise extending air diffusing body located between and spaced from said pair of panel members at the forward edges thereof, said air diusing body embodying a hollow portion and having a forwardly eX- tending wedge portion whereby it is adapted to effect two dimensional supersonic external compression of the air flow, a spanwise extending combustion chamber located between said pair of panel members, said combustion chamber having apertured walls forming a faired rearward continuation of said air diffusing body and diverging rearwardly to form converging air passages with said pair of panel members, ignition and fuel distribution means disposed within said hollow portion of said air diiusing body, and nozzle means located between said pair of panel members and forming a rearward continuation of the walls of said combustion chamber.

References Cited in the le of this patent UNITED STATES PATENTS 1,725,914 Hallowell Aug. 27, 1929 2,486,967 Morrisson Nov. 1, 1949 2,510,645 McMahan June 6, 1950 2,595,999 Way et al. May 6, 1952 2,597,610 Berliner May 20, 1952 2,631,425 Nordfors e Mar. 17, 1953 2,670,601 Williams et al. Mar. 2, 1954 2,672,333 Rocheville Mar. 16, 1954 2,735,263 Charshaan Feb. 21, 1956 

